XFOIL Version 6.94 Calculated polar for: UAG 88-143/20 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 1.000 0.8638 0.01232 0.00777 -0.1642 0.7532 1.0000 2.000 0.9529 0.01191 0.00728 -0.1603 0.7431 1.0000 3.000 1.0478 0.01142 0.00676 -0.1575 0.7324 1.0000 3.500 1.0864 0.01133 0.00669 -0.1543 0.7248 1.0000 4.000 1.1419 0.01095 0.00630 -0.1544 0.7188 1.0000 4.500 1.1869 0.01076 0.00614 -0.1525 0.7103 1.0000 5.000 1.2237 0.01074 0.00616 -0.1489 0.7000 1.0000 5.500 1.2609 0.01074 0.00620 -0.1453 0.6886 1.0000 6.000 1.2776 0.01096 0.00647 -0.1375 0.6739 1.0000 6.500 1.2982 0.01121 0.00676 -0.1308 0.6573 1.0000 7.000 1.3171 0.01175 0.00734 -0.1241 0.6381 1.0000 7.500 1.3354 0.01253 0.00812 -0.1179 0.6132 1.0000 8.000 1.3375 0.01412 0.00959 -0.1095 0.5711 1.0000 8.500 1.3341 0.01633 0.01156 -0.1009 0.5175 1.0000 9.000 1.3204 0.01933 0.01421 -0.0913 0.4524 1.0000 9.500 1.3157 0.02227 0.01691 -0.0838 0.3959 1.0000 10.000 1.3057 0.02591 0.02020 -0.0763 0.3263 1.0000 11.000 1.2552 0.03649 0.02936 -0.0603 0.1065 1.0000 11.500 1.2271 0.04279 0.03499 -0.0532 0.0119 1.0000 12.000 1.2387 0.04633 0.03869 -0.0500 0.0079 1.0000 12.500 1.2510 0.04995 0.04252 -0.0473 0.0067 1.0000 13.000 1.2540 0.05455 0.04732 -0.0443 0.0051 1.0000 13.500 1.2552 0.05959 0.05256 -0.0416 0.0048 1.0000 14.000 1.2573 0.06472 0.05789 -0.0395 0.0049 1.0000 14.500 1.2536 0.07069 0.06404 -0.0374 0.0046 1.0000 15.000 1.2340 0.07833 0.07180 -0.0348 0.0041 1.0000 15.500 1.2421 0.08315 0.07678 -0.0339 0.0042 1.0000 16.000 1.2498 0.08739 0.08113 -0.0318 0.0041 1.0000 16.500 1.2618 0.09158 0.08550 -0.0307 0.0042 1.0000 17.000 1.2720 0.09624 0.09047 -0.0296 0.0044 1.0000 17.500 1.3068 0.09726 0.09186 -0.0249 0.0060 1.0000